This disclosure relates generally to laminated components and, more particularly, to replacing an aperture in a laminated component.
Composite laminated structures typically include one or more plies of compressed reinforcement fabric layers bonded together by a resin matrix, such as an epoxy. Many laminated components include apertures for fasteners such as rivets. In aerospace applications, the rivets are typically titanium.
Over time, the apertures in the composite laminated structures are prone to wear from vibration. The apertures may also sometimes be double drilled during manufacture. In either instance, the apertures are enlarged or otherwise misplaced which minimizes their effective receipt of the fastener.
Various techniques have been developed to replace deformed or misplaced apertures. For example, in some laminated components, the plies are peeled back, cut off, and replaced as a structural restoration. A new aperture is then machined into the laminated component. Although effective, this technique is relatively complex, expensive, and requires specific tooling and knowledge of parent component structure and/or design.
Another technique involves replacing the deformed aperture with a metal bushing. Although effective, this technique may result in an undesirable thermal coefficient of expansion mismatch with the laminated structure and/or create undesirable galvanic conditions at the mating interface. Further, conventional bushings require significant removal of currently intact material which may weaken the laminate.
Yet another technique involves application of epoxy resin to fill the space of the deformed aperture which is then machined to form a replacement aperture. Although effective, this technique may not be significantly robust in that voids are frequently observed which may result in premature failure of the repair material. Further, the epoxy resin matrix may not attain proper design requirements, e.g. strength to avoid fastener pull through.